The present invention relates to a stationary vane in a gas turbine. More specifically, the present invention relates to a gas turbine stationary vane having a serpentine cooling air flow path with enhanced cooling effectiveness.
A gas turbine employs a plurality of stationary vanes that are circumferentially arranged in rows in a turbine section. Since such vanes are exposed to the hot gas discharging from the combustion section, cooling of these vanes is of the utmost importance. Typically, cooling is accomplished by flowing cooling air through cavities formed inside the vane airfoil.
According to one approach, cooling of the vane airfoil is accomplished by incorporating one or more tubular inserts into each of the airfoil cavities so that passages surrounding the inserts are formed between the inserts and the walls of the airfoil. The inserts have a number of holes distributed around their periphery that distribute the cooling air around these passages.
According to another approach, each airfoil cavity includes a number of radially extending passages, typically three, forming a serpentine array. Cooling air, supplied to the vane outer shroud, enters the first passage and flows radially inward until it reaches the vane inner shroud. A first portion of the cooling air exits the vane through the inner shroud and enters a cavity located between adjacent rows of rotor discs. The cooling air in the cavity serves to cool the faces of the discs. A second portion of the cooling air reverses direction and flows radially outward through the second passage until it reaches the outer shroud, whereupon it changes direction again and flows radially inward through the third passage, eventually exiting the blade from the third passage through holes in the trailing edge of the airfoil.
Various methods have been tried to increase the effectiveness of the cooling air flowing through the serpentine passages. One such approach involves the use of fins extending from the walls that form the passages. The use of both fins that extend perpendicular to the direction of flow and fins that are angled to the direction of flow have been tried. However, the ability of such schemes to adequately cool the vane airfoils is impaired in gas turbines in which the airfoils have large a cross-sectional area since this reduces the velocity, and hence the heat transfer coefficient, of the cooling air flowing through the passages. The cooling ability of such schemes is also impaired when used in conjunction with higher pressure ratio compressors, since the cooling air bled from such compressors is at a relatively high temperature.
Moreover, as the cooling air absorbs heat from the vane airfoil it becomes hotter. Consequently, the cooling air may become too hot to cool the trailing edge of the airfoil by the time it reaches the last serpentine passage, especially if more than three such passages are utilized. Also, excessive heat up of the cooling air as a result of airfoil cooling may render the cooling air too hot to cool the cavity between the discs.
One potential solution to these problems is to dramatically increase the cooling air supplied to the airfoil, thereby increasing the flow rate of the cooling air flowing through the passages. However, such a large increase in cooling air flow is undesirable. Although such cooling air eventually enters the hot gas flowing through the turbine section, little useful work is obtained from the cooling air, since it was not subject to heat up in the combustion section. Thus, to achieve high efficiency, it is crucial that the use of cooling air be kept to a minimum.
It is therefore desirable to provide a cooling scheme that significantly increases the cooling effectiveness of the cooling air flowing through the airfoil of a stationary vane in a gas turbine. It is also desirable to prevent excessive heat-up of the portion of the cooling air used to cool the trailing edge portion of the vane airfoil, as well as the portion of the cooling air used to cool the rotor discs.